Axial-radial cooling slots on inner air seal

ABSTRACT

An air seal may comprise an annular ring defined by at least a proximal surface, a distal surface, an aft side and a forward side. A channel may be disposed in the forward side of the air seal and/or the aft side of the air seal and may extend between the proximal surface and the distal surface. An additional channel extending from at least one of the forward side or the aft side may be disposed in the distal surface. The channel and the additional channel may be circumferentially in line. The channels may define a flow path for direction cooling air from a proximal side of the air seal to a distal side of the air seal.

FIELD

This disclosure relates generally to gas turbine engines, and moreparticularly to air seal arrangements for turbine engines.

BACKGROUND

Gas turbine engines are known, and typically include a fan deliveringair into a compressor, and also outwardly of the compressor as bypassair. The air is compressed in the compressor and delivered downstreaminto a combustion section where it is mixed with fuel and ignited.Products of this combustion pass downstream to a turbine and overturbine rotors, driving the turbine rotors to rotate. The turbine rotorsin turn rotate the compressors and fan.

The turbine may include multiple rotatable turbine blade arraysseparated by multiple stationary vane arrays. The turbine blades arecoupled to a rotor disk assembly which is configured to rotate about anengine axis. Typically, an air seal is provided between an aft rotordisk and a forward rotor disk and radially inward from a stationaryvane. The air seal may experience thermal loading during operation ofthe gas turbine engine.

SUMMARY

An air seal for a gas turbine engine comprising an annular ring definedby at least a proximal surface, a distal surface, an aft side and aforward side, a radial channel disposed in the air seal, the radialchannel disposed in at least one of the forward side or the aft side,the radial channel extending between the proximal surface and the distalsurface, and an axial channel disposed in the distal surface, the axialchannel extending from at least one of the forward side or the aft sideand circumferentially in line with the radial channel.

In various embodiments, the radial channel may be disposed on the aftside and the axial channel extends from the aft side. The radial channelmay be disposed on the forward side and the axial channel extends fromthe forward side. The radial channel and the axial channel may beconfigured to direct a cooling air from a proximal side of the air sealto a distal side of the air seal for cooling the air seal. The air sealmay be configured to receive the cooling air from an aperture disposedin a rotor disk leg, the rotor disk leg being located radially inwardfrom the air seal. The air seal may be configured to be coupled betweena forward rotor disk and an aft rotor disk. The air seal may compriseknife edges extending from the distal surface, the knife edgesconfigured to interface with a proximal surface of a vane platform. Theair seal may comprise a nickel-based alloy. A cross-section area of theradial channel may be greater than a cross-section area of the axialchannel.

A gas turbine engine may comprise a compressor section, a combustorsection, a turbine section, an aft blade disk, a forward blade disk, andan air seal coupled between the aft blade disk and the forward bladedisk. The air seal may comprise an annular ring defined by at least aproximal surface, a distal surface, an aft side and a forward side, aradial channel disposed in at least one of the forward side or the aftside and extending between the proximal surface and the distal surface,and an axial channel disposed in the distal surface and extending fromat least one of the forward side or the aft side and circumferentiallyin line with the radial channel.

In various embodiments, the radial channel may be disposed on the aftside and the axial channel extends from the aft side. The radial channelmay be disposed on the forward side and the axial channel extends fromthe forward side. The radial channel and the axial channel may beconfigured to direct a cooling air from a proximal side of the air sealto a distal side of the air seal for cooling the air seal. The air sealmay be configured to receive the cooling air from an aperture disposedin a rotor disk leg, the rotor disk leg being located radially inwardfrom the air seal. The air seal may be configured to be coupled betweena forward rotor disk and an aft rotor disk. The air seal may compriseknife edges extending from the distal surface, the knife edgesconfigured to interface with a proximal surface of a vane platform. Theair seal comprises a nickel-based alloy.

A method of manufacturing an air seal for a gas turbine engine maycomprise forming a radial channel in at least one of a forward side oran aft side of the air seal, the radial channel extending between aproximal surface and a distal surface, forming an axial channel in adistal surface of the air seal, the axial channel extending from atleast one of the forward side or the aft side and circumferentially inline with the radial channel.

In various embodiments, the forming the radial channel may be performedby milling the at least one of the forward side or the aft side of theair seal. The forming the radial channel and the forming the axialchannel may provide the radial channel having a cross-section area whichis greater than a cross-section area of the axial channel.

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated otherwise.These features and elements as well as the operation thereof will becomemore apparent in light of the following description and the accompanyingdrawings. It should be understood, however, the following descriptionand drawings are intended to be exemplary in nature and non-limiting.

BRIEF DESCRIPTION OF THE DRAWINGS

Various embodiments are particularly pointed out and distinctly claimedin the concluding portion of the specification. Below is a summary ofthe drawing figures, wherein like numerals denote like elements andwherein:

FIG. 1 illustrates a side cutaway view of a turbine engine, inaccordance with various embodiments;

FIG. 2 illustrates a cross-section view of a turbine section of a gasturbine engine, in accordance with various embodiments;

FIG. 3A illustrates an axial view of an air seal, in accordance withvarious embodiments;

FIG. 3B illustrates a radial view of an air seal, in accordance withvarious embodiments;

FIG. 3C illustrates a close-up axial view of channels formed in the airseal, in accordance with various embodiments;

FIG. 3D illustrates a perspective, cross-section view of the air sealhaving channels formed into the air seal, in accordance with variousembodiments;

FIG. 3E illustrates an isolated cross-section view of the air seal, thecross section intersecting channels disposed in the forward side of theair seal, in accordance with various embodiments;

FIG. 3F illustrates an isolated cross-section view of the air seal, thecross section intersecting channels disposed in the aft side of the airseal, in accordance with various embodiments;

FIG. 4A illustrates a cross-section view of the air seal in an installedposition, the cross section intersecting channels disposed in theforward side of the air seal, in accordance with various embodiments;

FIG. 4B illustrates a close up view of the forward side of the air sealof FIG. 4A, in accordance with various embodiments;

FIG. 5A illustrates a cross-section view of the air seal in an installedposition, the cross section intersecting channels disposed in the aftside of the air seal, in accordance with various embodiments;

FIG. 5B illustrates a close up view of the aft side of the air seal ofFIG. 5A, in accordance with various embodiments; and

FIG. 6 illustrates a flow chart of a method for manufacturing an airseal for a gas turbine engine, in accordance with various embodiments.

Elements and steps in the figures are illustrated for simplicity andclarity and have not necessarily been rendered according to anyparticular sequence. For example, steps that may be performedconcurrently or in different order are illustrated in the figures tohelp to improve understanding of embodiments of the present disclosure.

DETAILED DESCRIPTION

The detailed description of exemplary embodiments herein makes referenceto the accompanying drawings, which show exemplary embodiments by way ofillustration. While these exemplary embodiments are described insufficient detail to enable those skilled in the art to practice thedisclosure, it should be understood that other embodiments may berealized and that logical changes and adaptations in design andconstruction may be made in accordance with this disclosure and theteachings herein. Thus, the detailed description herein is presented forpurposes of illustration only and not of limitation. The scope of thedisclosure is defined by the appended claims. Furthermore, any referenceto singular includes plural embodiments, and any reference to more thanone component or step may include a singular embodiment or step. Also,any reference to attached, fixed, connected or the like may includepermanent, removable, temporary, partial, full and/or any other possibleattachment option. Additionally, any reference to without contact (orsimilar phrases) may also include reduced contact or minimal contact.Surface shading lines may be used throughout the figures to denotedifferent parts but not necessarily to denote the same or differentmaterials. In some cases, reference coordinates may be specific to eachfigure.

As used herein, “distal” refers to the direction radially outward, orgenerally, away from the axis of rotation of a turbine engine. As usedherein, “proximal” refers to a direction radially inward, or generally,towards the axis of rotation of a turbine engine.

With reference to FIG. 1, an exemplary gas turbine engine 2 is provided,in accordance with various embodiments. Gas turbine engine 2 is atwo-spool turbofan that generally incorporates a fan section 4, acompressor section 6, a combustor section 8 and a turbine section 10.Vanes 51 may be disposed throughout the gas turbine engine 2.Alternative engines include, for example, an augmentor section amongother systems or features. In operation, fan section 4 drives air alonga bypass flow-path B while compressor section 6 drives air along a coreflow-path C for compression and communication into combustor section 8then expansion through turbine section 10. Although depicted as aturbofan gas turbine engine 2 herein, it should be understood that theconcepts described herein are not limited to use with turbofans as theteachings is applicable to other types of turbine engines includingthree-spool architectures. A gas turbine engine may comprise anindustrial gas turbine (IGT) or a geared aircraft engine, such as ageared turbofan, or non-geared aircraft engine, such as a turbofan, ormay comprise any gas turbine engine as desired.

Gas turbine engine 2 generally comprises a low speed spool 12 and a highspeed spool 14 mounted for rotation about an engine central longitudinalaxis A-A′ relative to an engine static structure 16 via several bearingsystems 18-1, 18-2, and 18-3. It should be understood that bearingsystems is alternatively or additionally provided at locations,including for example, bearing system 18-1, bearing system 18-2, andbearing system 18-3.

Low speed spool 12 generally comprises an inner shaft 20 thatinterconnects a fan 22, a low pressure compressor section 24, e.g., afirst compressor section, and a low pressure turbine section 26, e.g., asecond turbine section. Inner shaft 20 is connected to fan 22 through ageared architecture 28 that drives the fan 22 at a lower speed than lowspeed spool 12. Geared architecture 28 comprises a gear assembly 42enclosed within a gear housing 44. Gear assembly 42 couples the innershaft 20 to a rotating fan structure. High speed spool 14 comprises anouter shaft 80 that interconnects a high pressure compressor section 32,e.g., second compressor section, and high pressure turbine section 34,e.g., first turbine section. A combustor 36 is located between highpressure compressor section 32 and high pressure turbine section 34. Amid-turbine frame 38 of engine static structure 16 is located generallybetween high pressure turbine section 34 and low pressure turbinesection 26. Mid-turbine frame 38 supports one or more bearing systems18, such as 18-3, in turbine section 10. Inner shaft 20 and outer shaft80 are concentric and rotate via bearing systems 18 about the enginecentral longitudinal axis A-A′, which is collinear with theirlongitudinal axes. As used herein, a “high pressure” compressor orturbine experiences a higher pressure than a corresponding “lowpressure” compressor or turbine.

The core airflow C is compressed by low pressure compressor section 24then high pressure compressor section 32, mixed and burned with fuel incombustor 36, then expanded over high pressure turbine section 34 andlow pressure turbine section 26. Mid-turbine frame 38 includes surfacestructures 40, which are in the core airflow path. Turbines 26, 34rotationally drive the respective low speed spool 12 and high speedspool 14 in response to the expansion.

An engine 2 may comprise a rotor blade 68 and a stator vane 51. Rotorblades 68 and stator vanes 51 may be arranged circumferentially aboutthe engine central longitudinal axis A-A′.

With reference to FIG. 2, a cross-section view of turbine section 200 isillustrated, in accordance with various embodiments. Yz-axes areprovided for ease of illustration. Turbine section 200 may include airseal 210, forward rotor disk 230, aft rotor disk 240, and vane platform250. A stator vane 252 may extend from vane platform 250. Stator vane252 may be stationary. Air seal 210 may be coupled between forward rotordisk 230 and aft rotor disk 240. Air seal 210 may comprise knife edge211 and knife edge 212. Knife edge 211 and knife edge 212 may extendradially outward from air seal 210. Knife edge 211 may extend towardsvane platform 250. Knife edge 212 may extend towards vane platform 250.Air seal 210 may be defined by a distal surface 215, a proximal surface214, a forward side 216, and an aft side 217.

Aft rotor disk 240 may include a leg 244 extending from aft rotor disk240 towards forward rotor disk 230. Leg 244 may be coupled to forwardrotor disk 230. Leg 244 may comprise an aperture 246. A first cavity 202may be located radially outward from air seal 210. First cavity 202 maybe partially defined by distal surface 215. First cavity 202 maycomprise a first pressure P1 during operation of turbine section 200.For example, first cavity 202 may comprise a first pressure P1 duringtakeoff and cruise conditions. A second cavity 204 may be locatedradially inward from air seal 210. Second cavity 204 may be at leastpartially defined by proximal surface 214 and leg 244. Second cavity 204may comprise a second pressure P2 during operation of turbine section200. A third cavity 206 may be located radially inward from leg 244.Third cavity 206 may comprise a third pressure P3 during operation ofturbine section 200. Aperture 246 may be configured and sized such thatpressure P2 is tends to be greater than pressure P1. In this regard,cooling air may enter second cavity 204 from third cavity 206 viaaperture 246, as illustrated by arrow 291. Cooling air may be directedforward, as illustrated by arrow 292, and/or may be directed aft, asillustrated by arrow 294. Cooling air directed in the forward directionmay enter channels, as will be discussed in greater detail herein,formed into forward side 216 of air seal 210 and directed into firstcavity 202, as illustrated by arrow 293. Cooling air directed in the aftdirection may enter channels, as will be discussed in greater detailherein, formed into aft side 217 of air seal 210 and directed radiallyoutwards, as illustrated by arrow 295.

With respect to FIG. 3A through FIG. 5B, elements with like elementnumbering, as depicted in FIG. 2, are intended to be the same and willnot necessarily be repeated for the sake of clarity.

With combined reference to FIG. 3A and FIG. 3B, an axial view and aradial view of air seal 210, respectively, are illustrated, inaccordance with various embodiments. Xy-axes and yz-axes, respectively,are provided for ease of illustration. Air seal 210 may comprise anannular ring 302. Air seal 210 may comprise a centerline axis 390.Centerline axis 390 may be substantially concentric with engine centrallongitudinal axis A-A′ (see FIG. 1) in response to air seal 210 being inan installed position.

With reference to FIG. 3C, an axial view of the forward side 216 airseal 210 is illustrated, in accordance with various embodiments. Xy-axesare provided for ease of illustration. A radial channel 320 may bedisposed in forward side 216 of air seal 210. Radial channel 320 mayextend between proximal surface 214 and distal surface 215. Radialchannel 320 may extend in a radial direction (y-direction). An axialchannel 322 may be disposed in distal surface 215 of air seal 210. Axialchannel 322 may extend from forward side 216 of air seal 210. Axialchannel 322 may extend in an axial direction (z-direction). Axialchannel 322 may be circumferentially in line with radial channel 320 asillustrated in FIG. 3C. In this regard, radial channel 320 and axialchannel 322 may interface at edge 321.

With reference to FIG. 3D, a perspective view of air seal 210 isillustrated, in accordance with various embodiments. Xyz-axes areprovided for ease of illustration. A radial channel 324 may be disposedin aft side 217 of air seal 210. Radial channel 324 may extend betweenproximal surface 214 (see FIG. 2) and distal surface 215. Radial channel324 may extend in a radial direction (y-direction). An axial channel 326may be disposed in distal surface 215 of air seal 210. Axial channel 326may extend from aft side 217 of air seal 210. Axial channel 326 mayextend in an axial direction (z-direction). Axial channel 326 may becircumferentially in line with radial channel 324 as illustrated in FIG.3D. In this regard, radial channel 324 and axial channel 326 mayinterface at edge 323.

With reference to FIG. 3E, an isolated cross-section view, with thecross-section intersecting radial channel 320 and axial channel 322, ofair seal 210 is illustrated, in accordance with various embodiments.

With reference to FIG. 3F, an isolated cross-section, view with thecross-section intersecting radial channel 324 and axial channel 326, ofair seal 210 is illustrated, in accordance with various embodiments.

With combined reference to FIG. 3E and FIG. 3F, radial channel 320 andaxial channel 322 may be circumferentially offset from radial channel324 and axial channel 326. However, it is contemplated herein thatradial channel 320 and axial channel 322 may be circumferentially inline with radial channel 324 and axial channel 326. Further, a pluralityof channels 320 and channels 322 may be circumferentially spaced aboutcenterline axis 390 (see FIG. 3B) in forward side 216. Still further, aplurality of channels 324 and channels 326 may be circumferentiallyspaced about centerline axis 390 (see FIG. 3B) in aft side 217.

With combined reference to FIG. 4A and FIG. 4B, a cross-section view,with the cross section intersecting radial channel 320 and axial channel322, of air seal 210, forward rotor disk 230, and aft rotor disk 240 inan installed position is illustrated, in accordance with variousembodiments. As previously mentioned, cooling air may flow into radialchannel 320, as illustrated by arrow 481, and into axial channel 322, asillustrated by arrow 482, and exit radially outward from air seal 210.Said cooling air may provide cooling to knife edge 211 and knife edge212. In this regard, radial channel 320 and axial channel 322 may aid inpreventing thermal fatigue of knife edge 211 and knife edge 212. In thisregard, a cooling air flow path, as illustrated by arrow 481 and arrow482 may be defined by radial channel 320, axial channel 322, and forwardrotor disk 230.

In various embodiments, the cross-section area of radial channel 320, asmeasured in the xz-plane, may be greater than the cross-section area ofaxial channel 322, as measured in the xy-plane. In this regard, axialchannel 322 may meter the flow of cooling air through said cooling airflow path. Providing a greater cross-section area of radial channel 320than the cross-section area of axial channel 322 may prevent radialchannel 320 from metering the flow of cooling air through said coolingair flow path in response to air seal moving axially relative to forwardrotor disk 230. For example, as illustrated in FIG. 4B, it should beappreciated, that the cross-section area, as measured in the xz-plane,of said cooling air flow path at radial channel 320 may decrease inresponse to air seal 210 moving in the negative z-direction relative toforward rotor disk 230. In this regard, gap G may decrease in responseto air seal 210 moving in the negative z-direction relative to forwardrotor disk 230. However, the cross-section area of said cooling air flowpath may not change in response to said movement of air seal 210. Inthis regard, providing a greater cross-section area of radial channel320 than the cross-section area of axial channel 322 may ensure thataxial channel 322 meters the flow of cooling air through said coolingair flow path, independent of gap G.

With combined reference to FIG. 5A and FIG. 5B, a cross-section view,with the cross section intersecting radial channel 324 and axial channel326, of air seal 210, forward rotor disk 230, and aft rotor disk 240 inan installed position is illustrated, in accordance with variousembodiments. As previously mentioned, cooling air may flow into radialchannel 324, as illustrated by arrow 581, and into axial channel 326, asillustrated by arrow 582, and exit radially outward from air seal 210.Said cooling air may provide cooling to air seal 210 and/or aft rotordisk 240. In this regard, radial channel 324 and axial channel 326 mayaid in preventing thermal fatigue of air seal 210 and/or aft rotor disk240. In this regard, a cooling air flow path, as illustrated by arrow581 and arrow 582 may be defined by radial channel 324, axial channel326, and aft rotor disk 240.

In various embodiments, the cross-section area of radial channel 324, asmeasured in the xz-plane, may be greater than the cross-section area ofaxial channel 326, as measured in the xy-plane. Providing a greatercross-section area of radial channel 324 than the cross-section area ofaxial channel 326 may ensure that axial channel 326 meters the flow ofcooling air through said cooling air flow path.

With reference to FIG. 6, a method 600 of manufacturing an air seal fora gas turbine engine is provided, in accordance with variousembodiments. Method 600 includes forming a radial channel in at leastone of a forward side or an aft side of an air seal (step 610). Method600 includes forming an axial channel in a distal surface of the airseal (step 620).

With combined reference to FIG. 2 and FIG. 6, step 610 may includeforming radial channel 320 in forward side 216 of air seal 210. Step 610may include forming radial channel 324 in aft side 217 of air seal 210.Step 620 may include forming axial channel 322 in distal surface 215 ofair seal 210. Step 620 may include forming axial channel 326 in distalsurface 215 of air seal 210.

Radial channel 320, axial channel 322, radial channel 324, and/or axialchannel 326 may be formed via a milling process. For example a mill endmay be used to cut or grind away material to form the channels. However,radial channel 320, axial channel 322, radial channel 324, and/or axialchannel 326 may be formed via any suitable process including additivemanufacturing methods and subtractive manufacturing methods.

In various embodiments, air seal 210 may be made of metal or metalalloys. In various embodiments, air seal 210 is made of a nickelsuperalloy such as an austenitic nickel-chromium-based alloy such asthat sold under the trademark Inconel® which is available from SpecialMetals Corporation of New Hartford, N.Y., USA. Air seal 210 may be madeof the same material as forward rotor disk 230 and/or aft rotor disk240, or may be made of a different material from forward rotor disk 230and/or aft rotor disk 240.

Benefits, other advantages, and solutions to problems have beendescribed herein with regard to specific embodiments. Furthermore, theconnecting lines shown in the various figures contained herein areintended to represent exemplary functional relationships and/or physicalcouplings between the various elements. It should be noted that manyalternative or additional functional relationships or physicalconnections may be present in a practical system. However, the benefits,advantages, solutions to problems, and any elements that may cause anybenefit, advantage, or solution to occur or become more pronounced arenot to be construed as critical, required, or essential features orelements of the disclosure. The scope of the disclosure is accordinglyto be limited by nothing other than the appended claims, in whichreference to an element in the singular is not intended to mean “one andonly one” unless explicitly so stated, but rather “one or more.”Moreover, where a phrase similar to “at least one of A, B, or C” is usedin the claims, it is intended that the phrase be interpreted to meanthat A alone may be present in an embodiment, B alone may be present inan embodiment, C alone may be present in an embodiment, or that anycombination of the elements A, B and C may be present in a singleembodiment; for example, A and B, A and C, B and C, or A and B and C.Systems, methods and apparatus are provided herein. In the detaileddescription herein, references to “one embodiment”, “an embodiment”,“various embodiments”, etc., indicate that the embodiment described mayinclude a particular feature, structure, or characteristic, but everyembodiment may not necessarily include the particular feature,structure, or characteristic. Moreover, such phrases are not necessarilyreferring to the same embodiment. Further, when a particular feature,structure, or characteristic is described in connection with anembodiment, it is submitted that it is within the knowledge of oneskilled in the art to affect such feature, structure, or characteristicin connection with other embodiments whether or not explicitlydescribed. After reading the description, it will be apparent to oneskilled in the relevant art(s) how to implement the disclosure inalternative embodiments.

Furthermore, no element, component, or method step in the presentdisclosure is intended to be dedicated to the public regardless ofwhether the element, component, or method step is explicitly recited inthe claims. No claim element is intended to invoke 35 U.S.C. 112(f)unless the element is expressly recited using the phrase “means for.” Asused herein, the terms “comprises”, “comprising”, or any other variationthereof, are intended to cover a non-exclusive inclusion, such that aprocess, method, article, or apparatus that comprises a list of elementsdoes not include only those elements but may include other elements notexpressly listed or inherent to such process, method, article, orapparatus.

What is claimed is:
 1. An air seal for a gas turbine engine comprising:an annular ring defined by at least a proximal surface, a distalsurface, an aft side and a forward side; a radial channel disposed inthe air seal, the radial channel disposed in at least one of the forwardside or the aft side, the radial channel extending between the proximalsurface and the distal surface; and an axial channel disposed in thedistal surface, the axial channel extending from at least one of theforward side or the aft side and circumferentially in line with theradial channel.
 2. The air seal of claim 1, wherein the radial channelis disposed on the aft side and the axial channel extends from the aftside.
 3. The air seal of claim 1, wherein the radial channel is disposedon the forward side and the axial channel extends from the forward side.4. The air seal of claim 1, wherein the radial channel and the axialchannel are configured to direct a cooling air from a proximal side ofthe air seal to a distal side of the air seal for cooling the air seal.5. The air seal of claim 4, wherein the air seal is configured toreceive the cooling air from an aperture disposed in a rotor disk leg,the rotor disk leg being located radially inward from the air seal. 6.The air seal of claim 1, wherein the air seal is configured to becoupled between a forward rotor disk and an aft rotor disk.
 7. The airseal of claim 1, wherein the air seal comprises knife edges extendingfrom the distal surface, the knife edges configured to interface with aproximal surface of a vane platform.
 8. The air seal of claim 1, whereinthe air seal comprises a nickel-based alloy.
 9. The air seal of claim 1,wherein a cross-section area of the radial channel is greater than across-section area of the axial channel.
 10. A gas turbine enginecomprising: a compressor section; a combustor section; a turbinesection; an aft blade disk; a forward blade disk; and an air sealcoupled between the aft blade disk and the forward blade diskcomprising: an annular ring defined by at least a proximal surface, adistal surface, an aft side and a forward side; a radial channeldisposed in at least one of the forward side or the aft side andextending between the proximal surface and the distal surface; and anaxial channel disposed in the distal surface and extending from at leastone of the forward side or the aft side and circumferentially in linewith the radial channel.
 11. The gas turbine engine of claim 10, whereinthe radial channel is disposed on the aft side and the axial channelextends from the aft side.
 12. The gas turbine engine of claim 10,wherein the radial channel is disposed on the forward side and the axialchannel extends from the forward side.
 13. The gas turbine engine ofclaim 10, wherein the radial channel and the axial channel areconfigured to direct a cooling air from a proximal side of the air sealto a distal side of the air seal for cooling the air seal.
 14. The gasturbine engine of claim 13, wherein the air seal is configured toreceive the cooling air from an aperture disposed in a rotor disk leg,the rotor disk leg being located radially inward from the air seal. 15.The gas turbine engine of claim 10, wherein the air seal is configuredto be coupled between a forward rotor disk and an aft rotor disk. 16.The gas turbine engine of claim 10, wherein the air seal comprises knifeedges extending from the distal surface, the knife edges configured tointerface with a proximal surface of a vane platform.
 17. The gasturbine engine of claim 10, wherein the air seal comprises anickel-based alloy.
 18. A method of manufacturing an air seal for a gasturbine engine comprising: forming a radial channel in at least one of aforward side or an aft side of the air seal, the radial channelextending between a proximal surface and a distal surface; forming anaxial channel in a distal surface of the air seal, the axial channelextending from at least one of the forward side or the aft side andcircumferentially in line with the radial channel.
 19. The method ofclaim 18, wherein the forming the radial channel is performed by millingthe at least one of the forward side or the aft side of the air seal.20. The method of claim 18, wherein the forming the radial channel andthe forming the axial channel provides the radial channel having across-section area which is greater than a cross-section area of theaxial channel.